Method of manufacturing an air cooled vane with film cooling pocket construction

ABSTRACT

An air cooled stator vane for a gas turbine engine is fabricated by casting the suction side wall and pressure side wall of the airfoil into separate halves and joining both halves at complementary sides and ribs, and forming pockets with a slot at the end of each pocket for flowing a film of cooling air over the exterior wall and drilling holes in the end of the pocket remote from the slot to fluidly connect the pocket with cooling air. The method includes in another embodiment a sheet metal sheath with pockets and slots identical to the pockets and slots in the shell configuration formed over the shell. Another embodiment includes perforated inserts extending in the cavity formed by the shell.

The invention was made under a U.S. Government contract and theGovernment has rights herein.

This is a division of copending application Ser. No. 07/550,008, filedon Jul. 9, 1990.

CROSS REFERENCE

The subject matter of this application is related to the subject matterof commonly assigned U.S. patent application Ser. No. 07/550,003 filedon even date herewith and entitled "Cooled Vane".

TECHNICAL FIELD

This invention relates to gas turbine engines and more particularly tothe cooling aspects of the vane and other stator components.

BACKGROUND ART

The technical community working in gas turbine engine technology haveand are continually expending considerable effort to improve the coolingaspects of the engine's component parts, particularly in the turbinearea. Obviously, improving the effectiveness of the cooling air resultsin either utilizing less air for cooling or operating the engine athigher temperature. Either situation attributes to an improvement in theperformance of the engine.

It is axiomatic that notwithstanding the enormous results anddevelopment that has occurred over the years the state-of-the-art offilm cooling and convection techniques are not optimum.

Some of the problems that adversely affect the cooling aspectsparticularly in vanes are (1) the pressure ratio across all of the filmholes cannot be optimized and (2) in vanes that incorporate conventionalinserts, the static pressure downstream of the insert is substantiallyconstant. Essentially in item (1) above the holes that operate with lessthan optimum pressure drop fail to produce optimum film cooling and initem (2) above a constant stator pressure adversely affects internalconvection.

One of the techniques that has been used with some degree of success iscoating of the airfoil sections of the vanes with a well known thermalbarrier coating. However, a coated vane conventionally requires drillingthe film coating hole after the coating process by a laser whichtypically results in a cylindrical hole compromising the film coolingeffectiveness, thus consequently reducing the effectiveness of thecoating. Moreover, flow control through the laser hole is moredifficult, presenting additional problems to the engine designer.

We have found that we can obviate the problems noted above and improvethe cooling effectiveness by providing in the vane a plurality ofpockets that form metering slots on the airfoil surface together withjudiciously located holes associated with each pocket for feedingcooling air to the metering slots which, in turn, effectively coalescethe air into a film of cooling air that flows across the externalsurface of the vane. The passageway from these located holes to theinclined slots places the cooling air in indirect parallel flow heatexchange relation with the gas path.

It is contemplated within the scope of this invention that the vane befabricated from either a total casting process or a partial castingprocess where a structural inner shell is cast and a sheath formed fromsheet metal encapsulates the shell.

A vane constructed in accordance with this invention affords amongstother advantages the following:

1) Film cooling effectiveness is optimized.

2) The film cooling system can adapt to thermal barrier coatings and thelike without film cooling compromise.

3) Convection is optimized since flow can be metered locally toheat-transfer requirements and overall pressure ratio.

4) In the sheet metal design a repair procedure can be accommodatedwhere distressed panels can be replaced without scrapping the totalpart.

5) A pressure side or suction side panel of the sheet metal designedvane may be optimized for flow and film coverage.

6) Improved cooling is achieved with hole and slot sizes that are largeenough to minimize internal plugging.

7) In the sheet metal configuration flexibility of material choices forthe external shell is significantly increased.

8) In the fully cast configuration the vane can be cast in halves whichoffer the most versatility in terms of achieving desired cooling flowsand film blowing parameters.

SUMMARY OF THE INVENTION

An object of this invention is to provide for a gas turbine engineimproved cooling effectiveness for the engine's vanes and/or statorcomponents.

A feature of this invention is to provide side walls that define theairfoil section of a vane having a plurality of pockets each having ametered slot for flowing film cooling air on the outer surface of theside wall and having judiciously located holes discreetly feedingcooling air into said pockets from a central passageway in the vanecommunicating with a source of cooling air. The airfoil surface in oneembodiment is formed from sheet metal supported to an inner cast shelland in another embodiment the vane including the airfoil section isfully cast. Still another embodiment employs a double layer of stampedsheet metal forming a 2-layer inner configuration. And still anotherembodiment includes a fully cast vane including pockets with judiciouslylocated holes as previously described, but also including inserts havinga plurality of apertures for feeding cooling air from the centralpassageway to the judiciously located holes.

The foregoing and other features and advantages of the present inventionwill become more apparent from the following description andaccompanying drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a partial view in schematic of the combustor, 1st turbine andvane of a gas turbine engine exemplary of the prior art.

FIG. 2 is a sectional view taken along lines 2--2 of FIG. 1 of a priorart vane.

FIG. 3 is a sectional view of a vane made in accordance with thisinvention showing the details thereof.

FIG. 4 is a sectional view of the inventive vane disclosing one of thesteps in manufacturing.

FIG. 5 is an enlarged view showing a portion of the pressure surface ofthe airfoil section of the vane in FIG. 3.

FIG. 6 is a partial view of an enlarged section of one of the pockets inFIG. 5.

FIG. 7 is a sectional view of the airfoil section depicting anotherembodiment of a vane incorporating this invention.

FIG. 8 is a partial view of an enlarged section of the stamped sheetmetal including the pocket in FIG. 7.

FIG. 9 is a sectional view of another embodiment of a vane employingthis invention.

FIG. 10 is a partial view of an enlarged section of the stamped sheetmetal sheath including the pocket in FIG. 9.

FIG. 11 is an exploded view in perspective of the embodiment depicted inFIG. 7.

FIG. 12 is an exploded view identical to FIG. 11 but including theplatform.

BEST MODE FOR CARRYING OUT THE INVENTION

While in its preferred embodiment this invention is being utilized inthe stator vane of the first turbine of a gas turbine engine, it will beunderstood by those skilled in this technology that the invention can beemployed in other vanes and other static components without departingfrom the scope of this invention. Notwithstanding the fact that thepreferred embodiment is a fully cast vane utilizing inserts, thepartially cast embodiment or fully cast embodiment without inserts areall deemed to be within the scope of this invention.

The invention can perhaps be best understood by first having anunderstanding of the state-of-the-art vane exemplified by the prior artdisclosed in FIGS. 1 and 2. As shown the vane generally indicated byreference numeral 10 is disposed between the turbine 12 and burner 14.The vane 10 is cooled by routing cool air obtained from the engine'scompressor section (not shown) via the passageways 16 and 18 which isdefined by the outer annular case 20 and outer liner 22 and innerannular case 24 and inner annular burner liner 26. Inserts 28 and 30opened at its base distribute the cool air from passageways 16 and 18through a plurality of holes formed in the walls thereof to a pluralityof holes formed in the pressure surface, suction surface, trailing andleading edges. Typically, flow entering the insert or impingement tubecircuit 28 from passageway 18 exits the vane as film air through filmholes in the leading edge 32, the pressure surface 34 and the suctionsurface 36. Flow entering the insert or impingement tube circuit 30 frompassageway 16 exits the vane as film air through film holes in thepressure surface 34 and suction surface 36 and as dump flow throughholes in the trailing edge 38. Platforms 35 and 37 on the inner andoutside diameter serve to attach the vane to the engine's turbine andcombustor cases and are opened to the compressor air flow.

What has been described is conventional in available gas turbine enginessuch as the JT9D, PW2037, PW4000 and F100 family of engines manufacturedby Pratt and Whitney division of United Technologies Corporation, theassignee common with this patent application. For the sake ofconvenience and simplicity only that portion germane to the invention isdescribed herein, and for further details the above noted engines areincorporated herein by reference.

The preferred embodiment is shown in FIGS. 3, 4, 5 and 6 which basicallyis a fully cast vane divided into three distinct regions, namely, theleading edge, the trailing edge and the side wall panels. The fully castvane 50 is comprised of the pressure side wall 52, the suction side wall54, the trailing edge 56 and the leading edge 58. The vane may be castin two halves as shown in FIG. 4 and bonded together by any suitablemeans, such as by transient liquid phase which is a well known joiningprocess and then brazed to the platform in a precision die, also a Wellknown technique. The ends of rib portions 61 and 63 extending inwardlymate when assembled to form a structural rib to prevent the vane frombulging due to the aerodynamic and pressure loads. Each side wall, i.e.the pressure side wall 52 and suction side wall 54, are cast with aplurality of pockets 60 (see FIG. 5) that are judiciously locatedadjacent the outer surface. A metering slot 62 is formed at the end ofeach pocket for exiting film air adjacent the outer surface of the sidewalls. A plurality of holes 64 are drilled internally of the pocket andcommunicate with the central passages 66 and 68 formed in the vane. Theholes 64 are judiciously located so that cooling air impinges on theback side of the pressure side wall 52 and suction side wall 54, turnsand flows toward the trailing edge in the mini passage 70 and exits outof metering slot 62 and effectively produce a film of cooling air. Eachpocket may include a pedestal or pedestals 74 consistent with eachapplication to enhance heat transfer. As noted in FIG. 5, each row ofpockets 60 is arranged so that alternate rows are staggered. As noted,the upper row of pockets is slightly displaced relative to the lower rowof pockets, assuring that a solid sheet of film cooling air isdistributed over the airfoil surface.

The fully cast vane 50 may include inserts or impingement tubes 76 and78 similar to the impingement tubes shown in the prior art (FIGS. 1 and2). A plurality of holes 80 in the walls of the impingement tubes 76 and78 serve to feed the side wall holes 64 of the pockets with the coolingair from the compressor section (not shown).

As shown in FIG. 6 cool air from the impingement tube flows throughholes 80 to impinge on the back surfaces of the pressure side wall 52effectuating impingement cooling and convection. The air then flows intothe holes 64 to impinge on each side of the wall 84 defining the pocket60 to likewise maximize cooling effectiveness. The air then turns andflows leftwardly as viewed in FIG. 6 which is in the direction of thetrailing edge and then out of metering slot 62 for laying a film of coolair adjacent the outer surface of the side wall.

By virtue of this invention not only is the convection processmaximized, it is also possible to attain a maximum coverage of filmcooling air downstream of the metering slot that extends over thesurface of the airfoil. Conventional trip strips 86 may be included onthe back side of the slot trailing edge to enhance heat transfer if sodesired.

In this design, as seen in FIG. 3 the leading edge 58 and trailing edge56 are cooled utilizing conventional technique although in certainembodiments as will be understood from the description to follow, theside walls are fed with cool air directly from the central passage inthe vane.

The airfoil section of the fully cast vane 50 as noted in FIG. 6 can becoated with a thermal barrier coating (TBC) similar to that used on theprior art vane as shown by the overlay 90. Since the exit slot flow areais several times larger than the metering holes, the metered slots withthe coating process are tolerant to TBC use. The TBC build-up closes theslots but not enough to shift the metering from the internal holes.Since the flow of cooling air is not affected by the TBC, the coatingprocess doesn't adversely affect the film cooling. In particular, whenTBC is a design feature, the exit slots are oversized such that theapplication of the TBC coats down the exit slots for an optimum arearatio of the exit slots to the metering holes, hence, the coolant to gasvelocity ratio and film cooling effectiveness are optimized.

FIGS. 7, 8, 9, 10, 11 and 12 exemplify vanes incorporating thisinvention that are fabricated from a partially cast process and stampedsheet metal sheaths defining the side wall airfoil section. Similar tothe fully cast vane construction, the embodiments depicted in FIGS. 7and 9 which are fabricated with a single and double liner layerconfiguration, divide the cooling into three distinct regions; namelythe leading edge, the trailing edge and the sidewall panels. Also,similarly these configurations combine backside impingement cooling,convection, surface liner backside impingement and a diffusing channelor metering slot discharging the coolant into the airfoil boundary layerwith an optimum blowing parameter.

In the single layer embodiment depicted in FIG. 7, the inner shell 100which is a structural member is formed in a hollow body defining thecentral passageway 102 and the shape of the airfoil section. The leadingedge 104 with conventional cooling techniques are cast in the shell sameis true for the trailing edge 106 which also employs conventionalcooling techniques. The shell includes a plurality of impingement holes122 that flow cooling air from the central passageway 102 which, similarto the vanes described above, is in communication with engine'scompressor air (not shown) exposed to the inner and outer diameter ofthe vane through the platforms (see FIG. 12, one being shown). Theseplatforms used for attaching the vanes to the engine's inner cases arecast on the inner and outer diameter of the shell. The outer linerlayers defining the outer surface of the airfoil section is stamp formedout of sheet metal and is contiguous to the outer surfaces of the shell.The sheet metal has stamp formed therein a plurality of shaped dimplesdefining pockets 112 extending over a portion of the surface of the sidewalls 108 on both the pressure side 114 and suction side 116. Pockets112 terminate in a slot 120 that is dimensioned to meter cooling flow toprovide an optimum blowing parameter and obtain an optimized film ofcooling air that flows adjacent the surface of the airfoil. The drilledholes 122 formed in the shell lead cooling air from the centralpassageway 102 to impinge on the backside of the trailing edge of themetered slot 120 in pocket 112 to effectuate impingement cooling andoptimize convection as the cool air flows through the pocket to themetered slot 120.

In assembly as best seen in FIGS. 7, 11 and 12 the stamped sheet metalliners defining the pressure surface and suction surface are dimensionedto fit into the recess formed between the leading edge 104 and trailingedge 106 adjacent the radial lips 130 and 132 respectively, on thepressure side and 134 and 136 respectively on the suction side. And theinner and outer diameter of the pressure and suction airfoil section fitinto and are trapped in slots 137 and 138 respectively-formed on both ofthe platforms (one being shown) (see FIG. 12).

As shown in FIG. 8 the layer defined by the stamp formed sheet metal forthe pressure side 114 and suction side 116 may be coated as in the priorart construction by a suitable thermal barrier coating and like thefully cast vane described above because the slot and holes aresignificantly large the coating does not adversely affect the coolingaspects.

FIGS. 9, 10, 11 and 12 disclose a multi-layer configuration similar tothe construction described in connection with the FIG. 7 embodiment. Forthe sake of convenience like reference numerals will designate similarcomponents in each of the embodiments. The outer liner layer 114 and 116are configured similar to the outer liner layer of the single layerembodiment as is the inner structural shell 100. In this embodiment anintermediate liner layer 140 is stamp formed out of sheet metal with adimple 142 that is complementary to the dimple in the outer liner layerand directs cooling air through holes 144 to impinge on the backside ofthe outer liner layer.

The assembly procedure for assembling the multi-layer design is similarto the assembly of the single liner layer design depicted in FIG. 7.

Both the single and multi-layer sheet metal designs permit the use ofdissimilar material for the liner layers and shell. This allows thedesigner a great latitude in the selection of materials such that hightemperature resistant but low strength material (such as wroughtmaterials) can be used for the outer liner layer and a high strength lowtemperature resistant material can be used for the inner shell. Thesedesigns also lend themselves for simplified and less costly repair andreplacement practices. Namely, a liner malfunction can be repaired bysimply replacing the liner rather than the entire vane, as theheretofore practice would be for certain impairments.

Although this invention has been shown and described with respect todetailed embodiments thereof, it will be understood by those skilled inthe art that various changes in form and detail thereof may be madewithout departing from the spirit and scope of the claimed invention.

We claim:
 1. The method of manufacturing an air cooled stator vane for agas turbine engine including the steps of:casting a first half of anouter shell shaped in an airfoil with an outer curvature defining aleading edge, a trailing edge and a suction surface; forming in the stepof casting the first half a plurality of spaced pockets in the outershell in rows extending over the suction surface of the shell andforming slots with a slot being located at the end of each pocket andpassages in the outer shell with each passage for each pocket extendingin the direction of the leading edge from the trailing edge connectingwith each of the slots; forming in the step of casting the first half amating face at the leading edge and a mating face at the trailing edge;casting a second half of the outer shell shaped in an airfoilcomplementing the first half for forming the stator vane with an outercurvature defining a leading edge, a trailing edge and a pressuresurface; forming in the step of casting the second half a plurality ofpockets in the outer shell in rows extending over the pressure surfaceand forming slots with a slot being located at the end of each pocketand passages in the outer shell with each passage for each pocketextending in the direction of the leading edge from the trailing edgeconnecting with each slot; forming in the step of casting the secondhalf a mating surface at the leading edge and a mating surface at thetrailing edge; machining apertures through the cast outer shell from theinterior of the outer shell so that each aperture is formed in a portionof the outer shell and extends from the interior of the outer shell tothe pocket at a point remote from the slot; and joining said halves ofairfoil shaped outer shells at the respective mating surfaces.
 2. Thesteps of manufacturing as claimed in claim 1 including the step offorming in the steps of casting the first half and the second half ofthe outer shell axially extending ribs on the underside of the pressuresurface and the underside of the suction surface with each rib havingcomplementary mating surfaces and joining in the step of joining therespective mating surfaces.
 3. The steps of manufacturing as claimed inclaim 2 including the step of forming from a sheet metal blank a pair ofimpingement tube inserts complementing the space within the shelladjacent each side of the joined rib and inserting said inserts into therespective spaces.
 4. The steps of manufacturing as claimed in claim 3including the step of coating the outer shell with a thermal barriercoating material.